Ice crystal protection for a gas turbine engine

ABSTRACT

A gas turbine engine includes a fan mounted to rotate about a main longitudinal axis; an engine core, including a compressor, a combustor, and turbine coupled to the compressor through a shaft; and reduction gearbox; wherein the compressor includes a plurality of stages, each stage including a respective rotor and stator, a first stage of the plurality of stages being arranged at an inlet and including a first rotor with a plurality of blades; each blade extending chordwise from a leading edge to a trailing edges, and from root to tip for a span height, wherein 0% of the span height corresponds to the root and 100% of span height corresponds to tip; wherein the ratio of a maximum leading edge radius of curvature of the first rotor blades to a minimum leading edge radius of curvature of the first rotor blades is included between 2.2 and 3.5.

The present disclosure relates generally to gas turbine engines, andmore specifically to arrangements for ice crystal protection for a gasturbine engine, in particular for a compressor of a gas turbine engine.

Gas turbine engines are used to power aircraft, watercraft, powergenerators, and the like. A gas turbine engine generally comprises, inaxial flow series from front to aft, an air intake, a fan, one or morecompressors, a combustor, one or more turbines, and an exhaust nozzle.Air entering the air intake is accelerated by the fan to produce two airflows: a first air flow (core engine flow) through an inlet duct intothe compressor and a second air flow (bypass flow) which passes througha bypass duct to provide propulsive thrust. Air entering the compressoris compressed, mixed with fuel and then fed into the combustor, wherecombustion of the air/fuel mixture occurs. The high temperature and highenergy exhaust fluids are then fed to the turbine, where the energy ofthe fluids is converted to mechanical energy to drive the compressor inrotation by suitable interconnecting shaft.

The compressor may be a multi-stage compressor, wherein each compressorstage comprises in axial flow series a row of rotor blades and a row ofstator vanes. A radially inner end of the rotor blades are connected toa hub that define an inner annulus. A casing circumscribes the rotorblades and the stator vanes and defines an outer annulus. The rotorblades each have a root and an aerofoil portion with a root, a tip, atrailing edge and a leading edge.

When operating in ice forming conditions (either super-cooled water iceor high altitude ice crystals), ice can accrete on vanes arranged in thecore inlet upstream of the front of the compressor, for example on anEngine Section Stator (ESS) or core inlet stator, or on Inlet GuideVanes (IGVs), either static or variable. When ice is shed from thevanes, which may be owing to aerodynamic loading or vibration, the iceis ingested by the compressor rotor blade stage immediately downstream.Depending on the size of the ice shed from the vanes, the ingested icecan damage the compressor rotor blades of the first stages, for examplethe first, second and even third stage.

Ice crystals and super-cooled water can also be ingested directlythrough the fan and travel along the inlet duct towards the compressor,impacting and potentially damaging the rotor blades of the first stages.

The ice threat can be significantly higher for geared turbofanarchitectures, as the fan rotates relatively slow making less likelythat ice is centrifuged outward above the core into the bypass duct.Moreover, as the gas turbine engine gets bigger, the fan blade diameterincreases and the gap between fan blades gets bigger, giving more ofline-of-sight through the fan blades into the core. In other words, icethreat can increase considerably in medium- and large-size gas turbineengines.

Different approaches have been proposed to protect the engine againstice accretion. According to one of these approaches, anti-icing systemsmay be provided for the vanes to prevent ice build-up, and to melt anyice that accumulates. For example, the vanes may be provided withelectrical heaters, or relatively hot air bled from the compressor maybe directed towards the vanes. Both systems are complicated to implementand detrimental to efficiency.

Another known approach is that to make the rotor blade leading edgegenerally thicker, to withstand ice crystal impacts and reduce bladedeflection. In this respect, the radius of curvature of the rotor bladeat the leading edge is increased, but conventional design criteriaapplied to medium and large geared engines result in rotor blades notoptimised in terms of weight and aerodynamic performance and efficiency,or in poor ice protection.

There is therefore a need for an improved system for ice crystalprotection for a gas turbine engine.

According to a first aspect, there is provided a gas turbine enginecomprising: a fan mounted to rotate about a main longitudinal axis; anengine core, comprising in axial flow series a compressor, a combustor,and a turbine coupled to the compressor through a shaft; a reductiongearbox that receives an input from the shaft and outputs drive to thefan so as to drive the fan at a lower rotational speed than the shaft.The compressor comprises a plurality of stages, each stage comprising arespective rotor and stator, a first stage of said plurality of stagesbeing arranged at an inlet and comprising a first rotor with a pluralityof first rotor blades, each blade extending chordwise from a leadingedge to a trailing edge, and from a root to a tip for a span height H,wherein 0% of the span height H corresponds to the root and 100% of thespan height H corresponds to the tip. The ratio of a maximum leadingedge radius of curvature of the first rotor blades to a minimum leadingedge radius of curvature of the first rotor blades is comprised between2.2 and 3.5.

The ratio of the maximum leading edge radius of curvature to the minimumleading edge radius of curvature of the first rotor blades between 2.2and 3.5 is relatively high compared to the blades of the first rotor innon-geared gas turbine engines. The present inventors have found thatsuch a relatively high ratio is particularly effective against ice andat the same time does not penalise the blades of the first rotor interms of weight and aerodynamic performance and efficiency as in theknow blades.

In the know blades, as the ratio is relatively low, the minimum leadingedge radius of curvature is relatively high compared to the maximumleading edge radius of curvature, which means that the blades aregenerally relatively thicker, therefore heavier and less performing andefficient.

On the contrary, the present inventors have found that it is notnecessary to increase the leading edge radius of curvature along thewhole blade, but it suffices to increase the leading edge radius ofcurvature in specific areas only to achieve an effective ice protection,without major penalties in terms of weight, performance and efficiency.Such specific areas may vary depending on the geometry of the inletduct, but the ratio of the maximum leading edge radius of curvature tothe minimum leading edge radius of curvature stays the same.

In other words, rather than making the whole leading edge thicker andtherefore penalising blade performance and engine efficiency, iceprotection can be achieved by selectively increasing the leading edgeradius of curvature in a specific area where most likely ice would havethe most negative impact, leaving the leading edge in remaining areas ata relatively small radiuses of curvature.

In substance, the present inventors have found that if the ratio of themaximum leading edge radius of curvature of the first rotor blades tothe minimum leading edge radius of curvature of the first rotor bladesis less than 2.2 and greater than 3.5, the blades can achieve neithersatisfactory ice protection nor performance/efficiency at the same time.

Indeed, when the ratio is less than 2.2, if the minimum leading edgeradius of curvature is kept relatively small to optimise performance,then the maximum leading edge radius of curvature is too small toprovide adequate ice protection; on the contrary, if the maximum leadingedge radius of curvature is optimised for ice protection, so it isrelatively large, then the minimum leading edge radius of curvature getsrelatively too large with a negative impact on performance/efficiency.

Analogously, when the ratio is greater than 3.5, if the minimum leadingedge radius of curvature is kept relatively small to optimiseperformance, then the maximum leading edge radius of curvature getsunnecessarily too large with negative effects on weight andperformance/efficiency); however, the minimum leading edge radius ofcurvature cannot be decreased below a safety level, resulting in amaximum leading edge radius of curvature still unnecessarily too large.

The disclosure may apply to blades with different leading edgecross-section. In embodiments, the first rotor blades may feature acircular or elliptical leading edge cross-section.

In embodiments of the disclosure, the ratio of the maximum leading edgeradius of curvature of the first rotor blades to the minimum leadingedge radius of curvature of the first blades may be equal or greaterthan 2.3, for example equal or greater than 2.4, or equal or greaterthan 2.5, or equal or greater than 2.6, or equal or greater than 2.7.

The ratio of the maximum leading edge radius of curvature of the firstrotor blades to the minimum leading edge radius of curvature of thefirst blades may be equal or less than 3.4, for example equal or lessthan 3.3, or equal or less than 3.2, or equal or less than 3.1, or equalor less than 3.0.

The ratio of the maximum leading edge radius of curvature of the firstrotor blades to the minimum leading edge radius of curvature of thefirst blades may be comprised between 2.2 and 3.5, for example between2.2 and 3.3, or between 2.2 and 3.0, or between 2.3 and 3.5, or between2.3 and 3.3, or between 2.3 and 3.0, or between 2.4 and 3.5, or between2.4 and 3.3, or between 2.4 and 3.0, or between 2.5 and 3.5, or between2.5 and 3.

The blades may comprise an aerofoil portion and a root, and a spanwisedirection is a direction extending between the tip and the root of theblades, and a chordwise direction is a direction extending between theleading edge and the trailing edge of the blades.

In the present disclosure, upstream and downstream are with respect tothe air flow through the compressor. Moreover, front and rear is withrespect to the gas turbine engine, i.e. the fan being in the front andthe turbine being in the rear of the engine.

In some embodiments the minimum leading edge radius of curvature may bepositioned at between 0% and 50% of the span height H, for examplebetween 20% and 40% of the span height H, or between 20% and 35% of thespan height H, or between 25% and 35% of the span height H. In someembodiments the leading edge radius of curvature may be constant between0% and 50% of the span height H and equal to the minimum leading edgeradius of curvature. In other words, the leading edge radius ofcurvature may present a flat distribution between 0% and 50% of the spanheight H.

In some embodiments the maximum leading edge radius of curvature may bepositioned at at least 60% of the span height H, for example at at least70% of the span height H, or at at least 80% of the span height H. Insome embodiments the maximum leading edge radius of curvature may bepositioned at between 60% and 100% of the span height H, for example atbetween 70% and 100% of the span height H, or between 80% and 100% ofthe span height H.

In some embodiments the leading edge radius of curvature may vary atbetween 60% and 100% of the span height H and may be at least twice aslarge as the minimum leading edge radius of curvature.

In some embodiments the leading edge radius of curvature between 85% and100%, or between 90% and 100%, of the span height H may be constant. Forexample the leading edge radius of curvature between 85% and 100%, orbetween 90% and 100%, of the span height H may be constant and equal tothe maximum leading edge radius of curvature.

In some embodiments the ratio of the leading edge radius of curvature at0% of the span height II to the minimum leading edge radius of curvaturemay be equal or greater than 1, for example equal or greater than 1.10,or equal or greater than 1.15, or equal or greater than 1.20. In someembodiments the ratio of the leading edge radius of curvature at 0% ofthe span height H to the minimum leading edge radius of curvature may beequal or less than 1.50, for example equal or less than 1.45, or equalor less than 1.40, or equal or less than 1.35. In some embodiments theratio of the leading edge radius of curvature at 0% of the span height Hto the minimum leading edge radius of curvature may be comprised between1 and 1.50, for example between 1 and 1.40, or between 1 and 1.35, orbetween 1 and 1.30, or between 1 and 1.25, or between 1 and 1.20, orbetween 1.10 and 1.40, or between 1.10 and 1.35, or between 1.15 and1.50, or between 1.15 and 1.40, or between 1.15 and 1.35, or between1.20 and 1.50, or between 1.20 and 1.45, or between 1.20 and 1.40, orbetween 1.20 and 1.35.

In some embodiments a ratio of the maximum leading edge radius ofcurvature to the leading edge radius of curvature at 0% of the spanheight H may be equal or higher than 1.7, for example equal or greaterthan 1.8, or equal or greater than 1.9, or equal or greater than 2.0, orequal or greater than 2.1, or equal or greater than 2.2, or equal orgreater than 2.3. In some embodiments the ratio of the maximum leadingedge radius of curvature to the leading edge radius of curvature at 0%of the span height H may be equal or less than 3.2, for example equal orless than 3.1, or equal or less than 3.0, or equal or less than 2.9, orequal or less than 2.8, or equal or less than 2.7. In some embodimentsthe ratio of the maximum leading edge radius of curvature to the leadingedge radius of curvature at 0% of the span height H may be comprisedbetween 1.7 and 3.2, for example between 1.7 and 3.0, or between 1.7 and2.7, or between 2.0 and 3.2, or between 2.0 and 3.0, or between 2.0 and2.7, or between 2.3 and 3.2, or between 2.3 and 3.0 or between 2.3 and2.7.

In some embodiments, the leading edge radius of curvature may decreaselinearly from 80% span height to 40% span height, for example from 75%to 40% span height, or from 70% to 40% span height, or from 65% to 40%span height, or from 80% to 45% span height, or from 80% to 50% spanheight, or from 75% to 45% span height, or from 70% to 45% span height,or from 65% to 45% span height.

In some embodiments the leading edge radius of curvature may decreaselinearly from 80% span height to 55% span height and the leading edgeradius of curvature may be constant and equal to the minimum leadingedge radius of curvature between 50% span height and 0% span height.

As previously stated, aspects of some embodiments may be advantageousfor medium- and large-size gas turbine engines. In some embodiments thefan may have a diameter equal or greater than 240 cm, for example equalor greater than 300 cm, or equal or greater than 350 cm, and the ratioof the maximum leading edge radius of curvature to the fan diameter maybe equal or greater than 1.4×10⁻⁴, for example equal or greater than1.5×10⁻⁴, or equal or greater than 1.6×10⁻⁴, or equal or greater than1.7×10⁻⁴.

In some embodiments the fan may have a diameter equal or less than 390cm, for example equal or less than 370 cm, or equal or greater than 350cm, and the ratio of the maximum leading edge radius of curvature to thefan diameter may be equal or less than 3.6×10⁻⁴, for example equal orgreater than 3.1×10⁻⁴, or equal or greater than 2.6×10⁻⁴, or equal orgreater than 2.1×10⁻⁴.

In some embodiments the fan may have a diameter comprised between 240 cmand 400 cm, for example between 240 cm and 360 cm, or between 280 cm and400 cm, or between 280 cm and 360 cm, or between 320 cm and 400 cm, orbetween 320 cm and 360 cm, and the ratio of the maximum leading edgeradius of curvature to the fan diameter may be comprised between1.4×10⁻⁴ and 3.6×10⁻⁴, for example between 1.4×10⁻⁴ and 3.0×10⁻⁴, orbetween 1.5×10⁻⁴ and 3.6×10⁻⁴, or between 1.5×10⁻⁴ and 3.0×10⁻⁴, orbetween 1.5×10⁻⁴ and 2.5×10⁻⁴.

In some embodiments, the maximum leading edge radius of curvature may beequal or greater than 0.4 mm, for example equal or greater than 0.5 mm,or equal or greater than 0.6 mm. The maximum leading edge radius ofcurvature may be equal or less than 0.9 mm, for example equal or lessthan 0.8 mm, or equal or less than 0.7 mm. The maximum leading edgeradius of curvature may be comprised between 0.4 mm and 0.9 mm, forexample between 0.4 mm and 0.8 mm, or between 0.5 mm and 0.9 mm, orbetween 0.5 mm and 0.8 mm, or between 0.6 mm and 0.9 mm, or between 0.6mm and 0.8 mm.

In some embodiments the minimum leading edge radius of curvature may beequal or less than 0.40 mm, for example equal or less than 0.35 mm, orequal or less than 0.30 mm. The minimum leading edge radius of curvaturemay be equal or greater than 0.16 mm, for example equal or greater than0.18 mm, or equal or greater than 0.20 mm, or equal or greater than 0.22mm. The minimum leading edge radius of curvature may be comprisedbetween 0.16 mm and 0.40 mm, for example between 0.16 mm and 0.35 mm, orbetween 0.16 mm and 0.30 mm, or between 0.18 mm and 0.40 mm, or between0.18 mm and 0.35 mm, or between 0.18 mm and 0.30 mm, or between 0.20 mmand 0.40 mm, or between 0.20 mm and 0.35, or between 0.20 mm and 0.30mm, or 0.22 mm and 0.30 mm.

In some embodiments the fan may rotate at cruise conditions at a speedbetween 1300 rpm and 2000 rpm, for example between 1300 rpm and 1900rpm, or between 1400 rpm and 1800 rpm, or between 1500 rpm and 1700 rpm.

In some embodiments the fan may comprise 16 to 24 fan blades, forexample 16 to 22 fan blades, or 16 to 20 fan blades, or 18 to 22 fanblades.

The compressor may comprise two or more stages. For example, thecompressor may comprise three or four stages. The compressor maycomprise less than twelve stages, for example less than eleven, or tenstages.

In come embodiments, the compressor may comprise 2 to 8 stages.

The compressor may be an intermediate pressure compressor and the gasturbine engine may further comprise a high pressure compressordownstream of the intermediate pressure compressor.

The turbine may be an intermediate pressure turbine and the gas turbineengine may further comprise a high pressure turbine upstream of theintermediate pressure compressor.

The shaft may be a first shaft and the gas turbine engine may furthercomprise a second shaft coupling the high pressure turbine to the highpressure compressor.

As noted elsewhere herein, the present disclosure may relate to a gasturbine engine. Such a gas turbine engine may comprise an engine corecomprising a turbine, a combustor, a compressor, and a core shaftconnecting the turbine to the compressor. Such a gas turbine engine maycomprise a fan (having fan blades) located upstream of the engine core.

Arrangements of the present disclosure may be particularly, although notexclusively, beneficial for fans that are driven via a gearbox.Accordingly, the gas turbine engine may comprise a gearbox that receivesan input from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft. The input tothe gearbox may be directly from the core shaft, or indirectly from thecore shaft, for example via a spur shaft and/or gear. The core shaft mayrigidly connect the turbine and the compressor, such that the turbineand compressor rotate at the same speed (with the fan rotating at alower speed).

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only be the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the example above.

The gearbox may be a reduction gearbox (in that the output to the fan isa lower rotational rate than the input from the core shaft). Any type ofgearbox may be used. For example, the gearbox may be a “planetary” or“star” gearbox, as described in more detail elsewhere herein.

The gearbox may have any desired reduction ratio (defined as therotational speed of the input shaft divided by the rotational speed ofthe output shaft), for example greater than 2.5, for example in therange of from 3 to 4.2, or 3.2 to 3.8, for example on the order of or atleast 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. Thegear ratio may be, for example, between any two of the values in theprevious sentence. Purely by way of example, the gearbox may be a “star”gearbox having a ratio in the range of from 3.1 or 3.2 to 3.8. In somearrangements, the gear ratio may be outside these ranges.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan andcompressor(s). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, where a secondcompressor is provided. By way of further example, the flow at the exitto the combustor may be provided to the inlet of the second turbine,where a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 0.28 to 0.32. These ratios maycommonly be referred to as the hub-to-tip ratio. The radius at the huband the radius at the tip may both be measured at the leading edge (oraxially forwardmost) part of the blade. The hub-to-tip ratio refers, ofcourse, to the gas-washed portion of the fan blade, i.e. the portionradially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches),260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm(around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350cm, 360 cm (around 140 inches), 370 cm (around 145 inches), 380 (around150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm (around 160inches) or 420 cm (around 165 inches). The fan diameter may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds), for examplein the range of from 240 cm to 280 cm or 330 cm to 380 cm.

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cmto 270 cm) may be in the range of from 1700 rpm to 2500 rpm, for examplein the range of from 1800 rpm to 2300 rpm, for example in the range offrom 1900 rpm to 2100 rpm. Purely by way of further non-limitativeexample, the rotational speed of the fan at cruise conditions for anengine having a fan diameter in the range of from 330 cm to 380 cm maybe in the range of from 1200 rpm to 2000 rpm, for example in the rangeof from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpmto 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 13 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tip) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.28, 0.29,0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (allunits in this paragraph being Jkg⁻¹ K⁻¹/(ms⁻¹)²). The fan tip loadingmay be in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds), forexample in the range of from 0.28 to 0.31 or 0.29 to 0.3.

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratiomay be in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds), forexample in the range of from 13 to 16, or 13 to 15, or 13 to 14. Thebypass duct may be substantially annular. The bypass duct may beradially outside the core engine. The radially outer surface of thebypass duct may be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressureupstream of the fan to the stagnation pressure at the exit of thehighest pressure compressor (before entry into the combustor). By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds), for example in the range of from 50 to 70.

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹'s, 105 Nkg⁻¹ s, 100 Nkg⁻¹ s, 95 Nkg⁻¹ s, 90 Nkg⁻¹ s, 85 Nkg⁻¹'s or80 Nkg⁻¹s. The specific thrust may be in an inclusive range bounded byany two of the values in the previous sentence (i.e. the values may formupper or lower bounds), for example in the range of from 80 Nkg⁻¹ s to100 Nkg⁻¹s, or 85 Nkg⁻¹ s to 95 Nkg⁻¹ s. Such engines may beparticularly efficient in comparison with conventional gas turbineengines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). Purely by way of example, agas turbine as described and/or claimed herein may be capable ofproducing a maximum thrust in the range of from 330 kN to 420 kN, forexample 350 kN to 400 kN. The thrust referred to above may be themaximum net thrust at standard atmospheric conditions at sea level plus15 degrees C. (ambient pressure 101.3 kPa, temperature 30 degrees C.),with the engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400K, 1450K, 1500K,1550K, 1600K or 1650K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 1800K to 1950K. The maximumTET may occur, for example, at a high thrust condition, for example at amaximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a bladed disc or a bladed ring. Any suitable methodmay be used to manufacture such a bladed disc or bladed ring. Forexample, at least a part of the fan blades may be machined from a blockand/or at least part of the fan blades may be attached to the hub/discby welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26fan blades.

As used herein, cruise conditions have the conventional meaning andwould be readily understood by the skilled person. Thus, for a given gasturbine engine for an aircraft, the skilled person would immediatelyrecognise cruise conditions to mean the operating point of the engine atmid-cruise of a given mission (which may be referred to in the industryas the “economic mission”) of an aircraft to which the gas turbineengine is designed to be attached. In this regard, mid-cruise is thepoint in an aircraft flight cycle at which 50% of the total fuel that isburned between top of climb and start of descent has been burned (whichmay be approximated by the midpoint—in terms of time and/ordistance—between top of climb and start of descent. Cruise conditionsthus define an operating point of, the gas turbine engine that providesa thrust that would ensure steady state operation (i.e. maintaining aconstant altitude and constant Mach Number) at mid-cruise of an aircraftto which it is designed to be attached, taking into account the numberof engines provided to that aircraft. For example where an engine isdesigned to be attached to an aircraft that has two engines of the sametype, at cruise conditions the engine provides half of the total thrustthat would be required for steady state operation of that aircraft atmid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruiseconditions are defined as the operating point of the engine thatprovides a specified thrust (required to provide—in combination with anyother engines on the aircraft—steady state operation of the aircraft towhich it is designed to be attached at a given mid-cruise Mach Number)at the mid-cruise atmospheric conditions (defined by the InternationalStandard Atmosphere according to ISO 2533 at the mid-cruise altitude).For any given gas turbine engine for an aircraft, the mid-cruise thrust,atmospheric conditions and Mach Number are known, and thus the operatingpoint of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be part of the cruise condition.For some aircraft, the cruise conditions may be outside these ranges,for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions (according to the International StandardAtmosphere, ISA) at an altitude that is in the range of from 10000 m to15000 m, for example in the range of from 10000 m to 12000 m, forexample in the range of from 10400 m to 11600 m (around 38000 ft), forexample in the range of from 10500 m to 1500 m, for example in the rangeof from 10600 m to 11400 m, for example in the range of from 10700 m(around 35000 ft) to 11300 m, for example in the range of from 10800 mto 11200 m, for example in the range of from 10900 m to 11100 m, forexample on the order of 11000 m. The cruise conditions may correspond tostandard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to anoperating point of the engine that provides a known required thrustlevel (for example a value in the range of from 30 kN to 35 kN) at aforward Mach number of 0.8 and standard atmospheric conditions(according to the International Standard Atmosphere) at an altitude of38000 ft (11582 m).

Purely by way of further example, the cruise conditions may correspondto an operating point of the engine that provides a known requiredthrust level (for example a value in the range of from 50 kN to 65 kN)at a forward Mach number of 0.85 and standard atmospheric conditions(according to the International Standard Atmosphere) at an altitude of35000 ft (10668 m).

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

According to an aspect, there is provided an aircraft comprising a gasturbine engine as described and/or claimed herein. The aircraftaccording to this aspect is the aircraft for which the gas turbineengine has been designed to be attached. Accordingly, the cruiseconditions according to this aspect correspond to the mid-cruise of theaircraft, as defined elsewhere herein.

According to an aspect, there is provided a method of operating a gasturbine engine as described and/or claimed herein. The operation may beat the cruise conditions as defined elsewhere herein (for example interms of the thrust, atmospheric conditions and Mach Number).

According to an aspect, there is provided a method of operating anaircraft comprising a gas turbine engine as described and/or claimedherein. The operation according to this aspect may include (or may be)operation at the mid-cruise of the aircraft, as defined elsewhereherein.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine;

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine;

FIG. 4 shows a first rotor blade of a compressor,

FIG. 5 is a partial schematic view, in cross-section, of the first rotorblade of FIG. 4 illustrating the difference between a maximum and aminimum leading edge radius of curvature;

FIG. 6 is a partial schematic view, in cross-section, of the first rotorblade of FIG. 4 illustrating the difference between the maximum leadingedge radius of curvature and the leading edge radius of curvature at 0%span height;

FIG. 7 is a partial schematic view, in cross-section, of the first rotorblade of FIG. 4 illustrating the difference between the minimum leadingedge radius of curvature and the leading edge radius of curvature at 0%span height.

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, a lowpressure compressor 14, a high pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low pressure turbine 19 anda core exhaust nozzle 20.

The low pressure compressor 14 and the high pressure compressor 15comprise respective pluralities of compressor stages, each stagecomprising a rotor and a stator. FIG. 2 shows a first stage 42 and asecond stage 44 of the low pressure compressor 14. The first stage 42 isarranged upstream of the second stage 44. The first stage 42 comprises afirst rotor with a row of first rotor blades 50 and, downstream thereof,a first stator with a row of first stator vanes 52. Although the lowpressure compressor 14 has been illustrated as comprising two stages, asnoted elsewhere herein, the low pressure compressor 14 may comprise adifferent number of stages, for example two to eight stages.

A nacelle 21 surrounds the gas turbine engine 10 and defines a bypassduct 22 and a bypass exhaust nozzle 18. The bypass airflow B flowsthrough the bypass duct 22. The fan 23 is attached to and driven by thelow pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is combusted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives the high pressure compressor 15 by asuitable interconnecting shaft 27. The fan 23 generally provides themajority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26,which is coupled to a sun wheel, or sun gear, 28 of the epicyclic geararrangement 30. Radially outwardly of the sun gear 28 and intermeshingtherewith is a plurality of planet gears 32 that are coupled together bya planet carrier 34. The planet carrier 34 constrains the planet gears32 to precess around the sun gear 28 in synchronicity whilst enablingeach planet gear 32 to rotate about its own axis. The planet carrier 34is coupled via linkages 36 to the fan 23 in order to drive its rotationabout the engine axis 9. Radially outwardly of the planet gears 32 andintermeshing therewith is an annulus or ring gear 38 that is coupled,via linkages 40, to a stationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3. There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2. For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2.

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 18, 20meaning that the flow through the bypass duct 22 has its own nozzle 18that is separate to and radially outside the core engine nozzle 20.However, this is not limiting, and any aspect of the present disclosuremay also apply to engines in which the flow through the bypass duct 22and the flow through the core 11 are mixed, or combined, before (orupstream of) a single nozzle, which may be referred to as a mixed flownozzle. One or both nozzles (whether mixed or split flow) may have afixed or variable area.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

FIG. 4 illustrates an exemplary first rotor blade 50 comprising a root54 and an aerofoil portion 56. The root 54 may have any suitable shapeadapted to be mounted on a disc (not illustrated). The aerofoil portion56 comprises a tip 58, opposite to the root 54, and a leading edge 60and a trailing edge 62. The aerofoil portion 56 further comprises apressure surface wall 64 extending from the leading edge 60 to thetrailing edge 62 and a suction surface wall 66 extending from theleading edge 60 to the trailing edge 62.

The aerofoil portion 56 extends along a spanwise direction S between theroot 54 and the tip 58 for a span height H, and along a chordwisedirection C between the leading edge 60 and the trailing edge 62.

The first rotor blade 50 may have a span height H comprised between 140mm and 220 mm and a true chord comprised between 80 mm and 160 mm.

The leading edge 60 features a leading edge radius of curvature variablealong the spanwise direction S. In other words, from the root 54, whichmay be considered at 0% of the span height H, to the tip 58, which maybe considered at 100% of the span height H, the leading edge radius ofcurvature varies as it will now be described in more detail withreference to FIGS. 5-7.

FIG. 5 shows two different transversal sections of the first rotor blade50 showing the radius of curvature of the leading edge 60 at differentspan heights. In detail, FIG. 5 shows a first section S1 taken alongline L1-L1 of FIG. 4 and containing the maximum leading edge radius ofcurvature R_(MAX) and a second section S2 taken along line L2-L2 of FIG.4 and containing the minimum leading edge radius of curvature R_(min).

It is to be noted that the leading edge 60 may not lay on a singleradial direction and the leading edge 60 at the first section S1 and atsecond section S2 may not be aligned along the same radial direction;thus the leading edge 60 at the first section S1 and at second sectionS2 are illustrated in FIG. 5 as coincident for sake of clarity only. Inother words, the blade 50 may have any suitable shape and the leadingedge 60 may extend along any suitable direction.

The first section S1 is taken at a span height H_(L1) corresponding to90% of the span height H. In other words, the maximum leading edgeradius of curvature R_(MAX) is arranged at 90% of the span height H. Inother non-illustrated embodiments the maximum leading edge radius ofcurvature R_(MAX) may be arranged at different span heights, for examplein a range between 60% and 100%, or 80% and 100% of the span height H.

The second section S2 is taken at a span height H_(1,2) corresponding to30% of the span height H. In other words, the minimum leading edgeradius of curvature R_(min), is arranged at 30% of the span height H. Inother non-illustrated embodiments the minimum leading edge radius ofcurvature R_(min) may be arranged at different span heights, for examplein a range between 20% and 40% of the span height H.

The ratio of the maximum leading edge radius of curvature R_(MAX) to theminimum leading edge radius of curvature R_(min) may be equal or greaterthan 2.2. Moreover, the ratio of the maximum leading edge radius ofcurvature R_(MAX) to the minimum leading edge radius of curvatureR_(min) may be equal or less than 3.5. In an embodiment, the ratio ofthe maximum leading edge radius of curvature R_(MAX) to the minimumleading edge radius of curvature R_(min) may be 3.0.

FIG. 6 shows the first section S1 of FIG. 5 containing the maximumleading edge radius of curvature R_(MAX) and a third section S3 takenalong line L3-L3 of FIG. 4 at a span height H_(L3) corresponding to 0%of the span height H. The leading edge radius of curvature at the spanheight H_(L3) is R_(0%).

As in FIG. 5, the maximum leading edge radius of curvature R_(MAX) andthe leading edge radius of curvature R_(0 %) may not necessarily bealigned along one radial direction and are illustrated as coincident forsake of clarity only.

The ratio of the maximum leading edge radius of curvature R_(MAX) to theleading edge radius of curvature R_(0%) may be equal or greater than1.7. Moreover, the ratio of maximum leading edge radius of curvatureR_(MAX) to the minimum leading edge radius of curvature R_(min) may beless than 3.2. In an embodiment, the ratio of maximum leading edgeradius of curvature R_(MAX) to the minimum leading edge radius ofcurvature R_(min) may be 2.4.

FIG. 7 shows the second section S2 containing the minimum leading edgeradius of curvature R_(min), and the third section S3 containing theleading edge radius of curvature R_(0%) at a span height of 0%.

The ratio of the leading edge radius of curvature R_(0%) to the minimumleading edge radius of curvature R_(min) may be equal or greater than1.1. Moreover, the ratio of the leading edge radius of curvature R_(0%)to the minimum leading edge radius of curvature R_(min) in may be lessthan 1.4. In an embodiment, the ratio of the leading edge radius ofcurvature R_(0%) to the minimum leading edge radius of curvature R_(min)may be 1.25.

As illustrated, the leading edge radius of curvature varies along thespan, decreasing from a maximum value at 80%-100% span height to aminimum value at 20%-40% span height, and then increasing again from theminimum value to the 0% span height value.

In non-illustrated embodiments the leading edge radius of curvatureR_(0%) may be equal to the minimum leading edge radius of curvatureR_(min), or in other words their ratio may be equal to 1. For example,the leading edge radius of curvature may be constant and equal to theminimum leading edge radius of curvature R_(min) between 0% and 50% ofthe span height H.

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

1. A gas turbine engine comprising: a fan mounted to rotate about a mainlongitudinal axis; an engine core, comprising in axial flow series acompressor, a combustor, and a turbine coupled to the compressor througha shaft; a reduction gearbox that receives an input from the shaft andoutputs drive to the fan so as to drive the fan at a lower rotationalspeed than the shaft; wherein the compressor comprises a plurality ofstages, each stage comprising a respective rotor and stator, a firststage of said plurality of stages being arranged at an inlet andcomprising a first rotor with a plurality of first rotor blades, eachblade extending chordwise from a leading edge to a trailing edges, andfrom a root to a tip for a span height H, wherein 0% of the span heightH corresponds to the root and 100% of the span height H corresponds tothe tip; wherein the ratio of a maximum leading edge radius of curvatureof the first rotor blades to a minimum leading edge radius of curvatureof the first rotor blades is comprised between 2.2 and 3.5.
 2. The gasturbine engine in accordance to claim 1, wherein the minimum leadingedge radius of curvature is positioned at between 0% and 50% of the spanheight H.
 3. The gas turbine engine in accordance to claim 1, whereinthe leading edge radius of curvature is constant between 0% and 50% ofthe span height H and equal to the minimum leading edge radius ofcurvature.
 4. The gas turbine engine in accordance to claim 1, whereinthe maximum leading edge radius of curvature is positioned at at least60% of the span height H.
 5. The gas turbine engine in accordance toclaim 1, wherein the ratio of the leading edge radius of curvature at 0%span height to the minimum leading edge radius of curvature is comprisedbetween 1 and 1.50.
 6. The gas turbine engine in accordance to claim 1,wherein the ratio of the maximum leading edge radius of curvature to theleading edge radius of curvature at 0% of the span height H is comprisedbetween 1.7 and 3.2.
 7. The gas turbine engine in accordance to claim 1,wherein the fan has a diameter comprised between 240 cm and 400 cm, andthe ratio of the maximum leading edge radius of curvature to the fandiameter is comprised between 1.4×10⁻⁴ and 3.6×10⁻⁴.
 8. The gas turbineengine in accordance to claim 1, wherein the maximum leading edge radiusof curvature is comprised between 0.4 mm and 0.9 mm.
 9. The gas turbineengine in accordance to claim 1, wherein the minimum leading edge radiusof curvature is comprised between 0.16 mm and 0.40 mm.
 10. The gasturbine engine in accordance to claim 1, wherein the first rotor bladesfeature a circular or elliptical leading edge cross-section.
 11. The gasturbine engine in accordance to claim 1, wherein the fan rotates at aspeed between 1300 rpm and 2000 rpm at cruise conditions.
 12. The gasturbine engine in accordance claim 1, wherein the fan comprises 16 to 24fan blades.
 13. The gas turbine engine according to claim 1, wherein thecompressor comprises 2 to 8 stages.
 14. The gas turbine engine accordingto claim 1, wherein the compressor is an intermediate pressurecompressor, the gas turbine engine further comprising a high pressurecompressor downstream of the intermediate pressure compressor; theturbine is an intermediate pressure turbine, the gas turbine enginefurther comprising a high pressure turbine upstream of the intermediatepressure compressor; and the shaft is a first shaft, the gas turbineengine further comprising a second shaft coupling the high pressureturbine to the high pressure compressor.